Method for setting a gear ratio of a fan drive gear system of a gas turbine engine

ABSTRACT

In one exemplary embodiment, a gas turbine engine includes an engine centerline longitudinal axis. A fan section includes a fan with a plurality of fan blades. The fan has a low corrected fan tip speed less than 1400 ft/sec. A bypass ratio is greater than 13 and less than 20. A fan pressure ratio less than 1.38 at cruise conditions of 0.8 Mach and about 35,000 feet. A speed reduction device comprises a gear system with a gear ratio of at least 2.6 and less than or equal to 4.1. A low pressure turbine is in communication with a first shaft. A high pressure turbine is in communication a second shaft. The first shaft is in communication with the fan through the speed reduction device. The low pressure turbine includes at least three stages and no more than four stages. The high pressure turbine includes two stages.

CROSS-REFERENCE TO RELATED APPLICATIONS

This disclosure is a continuation of U.S. application Ser. No.15/875,656 filed Jan. 19, 2018, which is a continuation of U.S.application Ser. No. 14/705,577 filed May 6, 2015, which is acontinuation in part of PCT/US2013/061115 filed on Sep. 23, 2013, whichis a continuation of U.S. application Ser. No. 13/758,075 filed Feb. 4,2013, which is now U.S. Pat. No. 8,753,065 issued on Jun. 17, 2014.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularlyto a method for setting a gear ratio of a fan drive gear system of a gasturbine engine.

A gas turbine engine may include a fan section, a compressor section, acombustor section, and a turbine section. Air entering the compressorsection is compressed and delivered into the combustor section where itis mixed with fuel and ignited to generate a high-speed exhaust gasflow. The high-speed exhaust gas flow expands through the turbinesection to drive the compressor and the fan section. Among othervariations, the compressor section can include low and high pressurecompressors, and the turbine section can include low and high pressureturbines.

Typically, a high pressure turbine drives a high pressure compressorthrough an outer shaft to form a high spool, and a low pressure turbinedrives a low pressure compressor through an inner shaft to form a lowspool. The fan section may also be driven by the inner shaft. A directdrive gas turbine engine may include a fan section driven by the lowspool such that a low pressure compressor, low pressure turbine, and fansection rotate at a common speed in a common direction.

A speed reduction device, which may be a fan drive gear system or othermechanism, may be utilized to drive the fan section such that the fansection may rotate at a speed different than the turbine section. Thisallows for an overall increase in propulsive efficiency of the engine.In such engine architectures, a shaft driven by one of the turbinesections provides an input to the speed reduction device that drives thefan section at a reduced speed such that both the turbine section andthe fan section can rotate at closer to optimal speeds.

Although gas turbine engines utilizing speed change mechanisms aregenerally known to be capable of improved propulsive efficiency relativeto conventional engines, gas turbine engine manufacturers continue toseek further improvements to engine performance including improvementsto thermal, transfer and propulsive efficiencies.

SUMMARY

In one exemplary embodiment, a gas turbine engine includes an enginecenterline longitudinal axis. A fan section includes a fan with aplurality of fan blades and rotatable about the engine centerlinelongitudinal axis. The fan has a low corrected fan tip speed less than1400 ft/sec. The low corrected fan tip speed is an actual fan tip speedat an ambient temperature divided by [(Tram °R)/(518.7 °R)] 0.5. Trepresents the ambient temperature in degrees Rankine. A bypass ratio isgreater than 13 and less than 20. A fan pressure ratio less than 1.48.The fan pressure ratio is measured across a fan blade alone. A speedreduction device comprises a gear system with a gear ratio of at least2.6 and less than or equal to 4.1. There is a plurality of bearingsystems. A low pressure turbine is in communication with a first shaft.A high pressure turbine is in communication a second shaft. The firstshaft and second shaft are concentric and mounted via at least one ofthe plurality of bearing systems for rotation about the enginecenterline longitudinal axis. The first shaft is in communication withthe fan through the speed reduction device. The low pressure turbineincludes at least three stages and no more than four stages. The highpressure turbine includes two stages. The low pressure turbine includesat least one rotor constrained by a first stress level. At least one ofthe plurality of fan blades of the fan is constrained by a second stresslevel and has a fan tip speed boundary condition. The gear ratio isconfigured such that in operation the fan blade does not exceed the fantip speed boundary condition or the second stress level. The lowpressure turbine rotor does not exceed the first stress level.

In a further embodiment of any of the above, the low pressure turbineincludes four stages.

In a further embodiment of any of the above, the gear system is a stargear system with a ring gear and a sun gear. The gear ratio isdetermined by measuring a diameter of the ring gear and dividing thatdiameter by the diameter of the sun gear.

In a further embodiment of any of the above, the fan pressure ratio isless than 1.38.

In a further embodiment of any of the above, the fan pressure ratio isless than 1.38.

In another exemplary embodiment, a gas turbine engine includes an enginecenterline longitudinal axis. A fan section includes a fan with aplurality of fan blades and rotatable about the engine centerlinelongitudinal axis. The fan has a low corrected fan tip speed less than1400 ft/sec. The low corrected fan tip speed is an actual fan tip speedat an ambient temperature divided by [(Tram °R)/(518.7 °R)]^(0.5). Trepresents the ambient temperature in degrees Rankine. A bypass ratio isgreater than 11.0 and less than 22.0. A fan pressure ratio is less than1.38. The fan pressure ratio is measured across a fan blade alone. Aspeed reduction device comprises a gear system. There is a plurality ofbearing systems. A low pressure turbine is in communication with a firstshaft. A high pressure turbine is in communication with a second shaft.The first shaft and second shaft are concentric and mounted via at leastone of the plurality of bearing systems for rotation about the enginecenterline longitudinal axis. The first shaft is in communication withthe fan through the speed reduction device. The high pressure turbineincludes two stages. The low pressure turbine includes at least onerotor constrained by a first stress level. At least one of the pluralityof fan blades of the fan is constrained by a second stress level and hasa fan tip speed boundary condition. The gear system is configured suchthat in operation the fan blade does not exceed the fan tip speedboundary condition or the second stress level. The low pressure turbinerotor does not exceed the first stress level.

In a further embodiment of any of the above, the low pressure turbineincludes five stages.

In a further embodiment of any of the above, a low pressure compressorincluding three stages. The low pressure turbine drives the low pressurecompressor.

In a further embodiment of any of the above, the low pressure turbineincludes at least three stages and no more than four stages.

In a further embodiment of any of the above, the speed reduction deviceincludes a gear ratio less than or equal to 4.1.

In a further embodiment of any of the above, the gear ratio is greaterthan or equal to 2.6.

In a further embodiment of any of the above, the low pressure turbineincludes four stages.

In a further embodiment of any of the above, a mid-turbine frame isarranged between the high pressure turbine and the low pressure turbine.The mid-turbine frame supports at least one bearing system.

In a further embodiment of any of the above, the mid-turbine frameincludes one or more airfoils that extend in a flow path.

In a further embodiment of any of the above, a low pressure compressorincluding three stages. The low pressure turbine drives the low pressurecompressor.

In a further embodiment of any of the above, a high pressure compressorincludes eight stages. The high pressure turbine drives the highpressure compressor.

In a further embodiment of any of the above, the star gear systemfurther comprising five intermediate gears.

In a further embodiment of any of the above, the gear system is a stargear system with a ring gear, a sun gear, and a star gear ratio. Thestar gear ratio is determined by measuring a diameter of the ring gearand dividing that diameter by the diameter of the sun gear.

In a further embodiment of any of the above, a mid-turbine frame isarranged between the high pressure turbine and the low pressure turbine.The mid-turbine frame supports at least one bearing system and includesone or more airfoils.

In another exemplary embodiment, a gas turbine engine includes an enginecenterline longitudinal axis. A fan section includes a fan with aplurality of fan blades and rotatable about the engine centerlinelongitudinal axis and has a low corrected fan tip speed less than 1400ft/sec. The low corrected fan tip speed is an actual fan tip speed at anambient temperature divided by [(Tram °R)/(518.7 °R)]^(0.5). Trepresents the ambient temperature in degrees Rankine. A bypass ratio ofgreater than 11.0. A speed reduction device comprising a gear system.There is a plurality of bearing systems. A low pressure turbine is incommunication with a first shaft and includes a pressure ratio greaterthan about 5:1. The low pressure turbine includes an inlet that has aninlet pressure and an outlet having an outlet pressure. The pressureratio of the low pressure turbine is a ratio of the inlet pressure tothe outlet pressure. A high pressure turbine in communication with asecond shaft. The first shaft and second shaft are concentric andmounted via at least one of the plurality of bearing systems forrotation about the engine centerline longitudinal axis. The first shaftis in communication with the fan through the speed reduction device. Thelow pressure turbine includes at least three stages and no more thanfour stages. The low pressure turbine includes at least one rotorconstrained by a first stress level. At least one of the plurality offan blades of the fan is constrained by a second stress level and has afan tip speed boundary condition. The gear system is configured suchthat in operation the fan blade does not exceed the fan tip speedboundary condition or the second stress level. The low pressure turbinerotor does not exceed the first stress level.

In a further embodiment of any of the above, the high pressure turbineincludes two stages.

In a further embodiment of any of the above, the low pressure turbineincludes three stages.

In a further embodiment of any of the above, the gear system has a gearratio of less than or equal to 4.1.

In a further embodiment of any of the above, the gear reduction ratio isat least 2.6.

In a further embodiment of any of the above, the gear system is a stargear system with a ring gear and a sun gear. The gear ratio isdetermined by measuring a diameter of the ring gear and dividing thatdiameter by the diameter of the sun gear.

In a further embodiment of any of the above, the low pressure turbineincludes four stages.

In another exemplary embodiment, a gas turbine engine includes an enginecenterline longitudinal axis. A fan section includes a fan with aplurality of fan blades and rotatable about the engine centerlinelongitudinal axis and has a low corrected fan tip speed less than 1400ft/sec. The low corrected fan tip speed is an actual fan tip speed at anambient temperature divided by [(Tram °R)/(518.7 °R)]^(0.5). Trepresents the ambient temperature in degrees Rankine. A bypass ratio isgreater than 11 and less than 22. A speed reduction device comprises agear system with a gear ratio. A plurality of bearing systems. A lowpressure turbine is in communication with a first shaft and a highpressure turbine in communication with a second shaft. The first shaftand second shaft are concentric and mounted via at least one of theplurality of bearing systems for rotation about the engine centerlinelongitudinal axis. The first shaft is in communication with the fanthrough the speed reduction device. The low pressure turbine includesfour stages. The low pressure turbine includes at least one rotorconstrained by a first stress level. At least one of the plurality offan blades of the fan is constrained by a second stress level and has afan tip speed boundary condition. The gear ratio is configured such thatin operation the fan blade does not exceed the fan tip speed boundarycondition or the second stress level. The low pressure turbine rotordoes not exceed the first stress level.

In a further embodiment of any of the above, the high pressure turbineincludes two stages.

In a further embodiment of any of the above, the gear reduction ratio isat least 2.6 and less than or equal to 4.1.

In a further embodiment of any of the above, the low pressure turbinehas a pressure ratio greater than 5:1. The low pressure turbine includesan inlet that has an inlet pressure and an outlet that has an outletpressure. The pressure ratio of the low pressure turbine is a ratio ofthe inlet pressure to the outlet pressure.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic, cross-sectional view of an example gasturbine engine.

FIG. 2 illustrates a schematic view of one configuration of a low speedspool that can be incorporated into a gas turbine engine.

FIG. 3 illustrates a fan drive gear system that can be incorporated intoa gas turbine engine.

FIG. 4 shows another embodiment.

FIG. 5 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The exemplarygas turbine engine 20 is a two-spool turbofan engine that generallyincorporates a fan section 22, a compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engines might includean augmentor section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B, while thecompressor section 24 drives air along a core flow path C forcompression and communication into the combustor section 26. The hotcombustion gases generated in the combustor section 26 are expandedthrough the turbine section 28. Although depicted as a two-spoolturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto two-spool turbofan engines and these teachings could extend to othertypes of engines, including but not limited to, three-spool enginearchitectures.

The exemplary gas turbine engine 20 generally includes a low speed spool30 and a high speed spool 32 mounted for rotation about an enginecenterline longitudinal axis A. The low speed spool 30 and the highspeed spool 32 may be mounted relative to an engine static structure 33via several bearing systems 31. It should be understood that otherbearing systems 31 may alternatively or additionally be provided, andthe location of bearing systems 31 may be varied as appropriate to theapplication.

The low speed spool 30 generally includes an inner shaft 34 thatinterconnects a fan 36, a low pressure compressor 38 and a low pressureturbine 39. The inner shaft 34 can be connected to the fan 36 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 45, such as a fan drive gear system50 (see FIGS. 2 and 3). The speed change mechanism drives the fan 36 ata lower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 35 that interconnects a high pressure compressor37 and a high pressure turbine 40. In this embodiment, the inner shaft34 and the outer shaft 35 are supported at various axial locations bybearing systems 31 positioned within the engine static structure 33.

A combustor 42 is arranged in exemplary gas turbine 20 between the highpressure compressor 37 and the high pressure turbine 40. A mid-turbineframe 44 may be arranged generally between the high pressure turbine 40and the low pressure turbine 39. The mid-turbine frame 44 can supportone or more bearing systems 31 of the turbine section 28. Themid-turbine frame 44 may include one or more airfoils 46 that extendwithin the core flow path C. It will be appreciated that each of thepositions of the fan section 22, compressor section 24, combustorsection 26, turbine section 28, and fan drive gear system 50 may bevaried. For example, gear system 50 may be located aft of combustorsection 26 or even aft of turbine section 28, and fan section 22 may bepositioned forward or aft of the location of gear system 50.

The inner shaft 34 and the outer shaft 35 are concentric and rotate viathe bearing systems 31 about the engine centerline longitudinal axis A,which is co-linear with their longitudinal axes. The core airflow iscompressed by the low pressure compressor 38 and the high pressurecompressor 37, is mixed with fuel and burned in the combustor 42, and isthen expanded over the high pressure turbine 40 and the low pressureturbine 39. The high pressure turbine 40 and the low pressure turbine 39rotationally drive the respective high speed spool 32 and the low speedspool 30 in response to the expansion.

In a non-limiting embodiment, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 45can include an epicyclic gear train, such as a planetary gear system, astar gear system, or other gear system. The geared architecture 45enables operation of the low speed spool 30 at higher speeds, which canenable an increase in the operational efficiency of the low pressurecompressor 38 and low pressure turbine 39, and render increased pressurein a fewer number of stages.

The pressure ratio of the low pressure turbine 39 can be pressuremeasured prior to the inlet of the low pressure turbine 39 as related tothe pressure at the outlet of the low pressure turbine 39 and prior toan exhaust nozzle of the gas turbine engine 20. In one non-limitingembodiment, the bypass ratio of the gas turbine engine 20 is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor 38, and the low pressure turbine 39has a pressure ratio that is greater than about five (5:1). In anothernon-limiting embodiment, the bypass ratio is greater than 11 and lessthan 22, or greater than 13 and less than 20. It should be understood,however, that the above parameters are only exemplary of a gearedarchitecture engine or other engine using a speed change mechanism, andthat the present disclosure is applicable to other gas turbine engines,including direct drive turbofans. In one non-limiting embodiment, thelow pressure turbine 39 includes at least one stage and no more thaneight stages, or at least three stages and no more than six stages. Inanother non-limiting embodiment, the low pressure turbine 39 includes atleast three stages and no more than four stages.

In this embodiment of the exemplary gas turbine engine 20, a significantamount of thrust is provided by the bypass flow path B due to the highbypass ratio. The fan section 22 of the gas turbine engine 20 isdesigned for a particular flight condition—typically cruise at about 0.8Mach and about 35,000 feet. This flight condition, with the gas turbineengine 20 at its best fuel consumption, is also known as bucket cruiseThrust Specific Fuel Consumption (TSFC). TSFC is an industry standardparameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. In another non-limitingembodiment of the example gas turbine engine 20, the Fan Pressure Ratiois less than 1.38 and greater than 1.25. In another non-limitingembodiment, the fan pressure ratio is less than 1.48. In anothernon-limiting embodiment, the fan pressure ratio is less than 1.52. Inanother non-limiting embodiment, the fan pressure ratio is less than1.7. Low Corrected Fan Tip Speed is the actual fan tip speed divided byan industry standard temperature correction of [(Tram °R)/(518.7°R)]^(0.5), where T represents the ambient temperature in degreesRankine. The Low Corrected Fan Tip Speed according to one non-limitingembodiment of the example gas turbine engine 20 is less than about 1150fps (351 m/s). The Low Corrected Fan Tip Speed according to anothernon-limiting embodiment of the example gas turbine engine 20 is lessthan about 1400 fps (427 m/s). The Low Corrected Fan Tip Speed accordingto another non-limiting embodiment of the example gas turbine engine 20is greater than about 1000 fps (305 m/s).

FIG. 2 schematically illustrates the low speed spool 30 of the gasturbine engine 20. The low speed spool 30 includes the fan 36, the lowpressure compressor 38, and the low pressure turbine 39. The inner shaft34 interconnects the fan 36, the low pressure compressor 38, and the lowpressure turbine 39. The inner shaft 34 is connected to the fan 36through the fan drive gear system 50. In this embodiment, the fan drivegear system 50 provides for counter-rotation of the low pressure turbine39 and the fan 36. For example, the fan 36 rotates in a first directionD1, whereas the low pressure turbine 39 rotates in a second direction D2that is opposite of the first direction D1.

FIG. 3 illustrates one example embodiment of the fan drive gear system50 incorporated into the gas turbine engine 20 to provide forcounter-rotation of the fan 36 and the low pressure turbine 39. In thisembodiment, the fan drive gear system 50 includes a star gear systemwith a sun gear 52, a ring gear 54 disposed about the sun gear 52, and aplurality of star gears 56 having journal bearings 57 positioned betweenthe sun gear 52 and the ring gear 54. A fixed carrier 58 carries and isattached to each of the star gears 56. In this embodiment, the fixedcarrier 58 does not rotate and is connected to a grounded structure 55of the gas turbine engine 20.

The sun gear 52 receives an input from the low pressure turbine 39 (seeFIG. 2) and rotates in the first direction D1 thereby turning theplurality of star gears 56 in a second direction D2 that is opposite ofthe first direction D1. Movement of the plurality of star gears 56 istransmitted to the ring gear 54 which rotates in the second direction D2opposite from the first direction D1 of the sun gear 52. The ring gear54 is connected to the fan 36 for rotating the fan 36 (see FIG. 2) inthe second direction D2.

A star system gear ratio of the fan drive gear system 50 is determinedby measuring a diameter of the ring gear 54 and dividing that diameterby a diameter of the sun gear 52. In one embodiment, the star systemgear ratio of the geared architecture 45 is between 1.5 and 4.1. Inanother embodiment, the system gear ratio of the fan drive gear system50 is between 2.6 and 4.1. When the star system gear ratio is below 1.5,the sun gear 52 is relatively much larger than the star gears 56. Thissize differential reduces the load the star gears 56 are capable ofcarrying because of the reduction in size of the star gear journalbearings 57. When the star system gear ratio is above 4.1, the sun gear52 may be much smaller than the star gears 56. This size differentialincreases the size of the star gear 56 journal bearings 57 but reducesthe load the sun gear 52 is capable of carrying because of its reducedsize and number of teeth. Alternatively, roller bearings could be usedin place of journal bearings 57.

Improving performance of the gas turbine engine 20 begins by determiningfan tip speed boundary conditions for at least one fan blade of the fan36 to define the speed of the tip of the fan blade. The maximum fandiameter is determined based on the projected fuel burn derived frombalancing engine efficiency, mass of air through the bypass flow path B,and engine weight increase due to the size of the fan blades.

Boundary conditions are then determined for the rotor of each stage ofthe low pressure turbine 39 to define the speed of the rotor tip and todefine the size of the rotor and the number of stages in the lowpressure turbine 39 based on the efficiency of low pressure turbine 39and the low pressure compressor 38.

Constraints regarding stress levels in the rotor and the fan blade areutilized to determine if the rotary speed of the fan 36 and the lowpressure turbine 39 will meet a desired number of operating life cycles.If the stress levels in the rotor or the fan blade are too high, thegear ratio of the fan drive gear system 50 can be lowered and the numberof stages of the low pressure turbine 39 or annular area of the lowpressure turbine 39 can be increased.

FIG. 4 shows an embodiment 100, wherein there is a fan drive turbine 108driving a shaft 106 to in turn drive a fan rotor 102. A gear reduction104 may be positioned between the fan drive turbine 108 and the fanrotor 102. This gear reduction 104 may be structured and operate likethe geared architecture 45 disclosed above. A compressor rotor 110 isdriven by an intermediate pressure turbine 112, and a second stagecompressor rotor 114 is driven by a turbine rotor 116. A combustionsection 118 is positioned intermediate the compressor rotor 114 and theturbine section 116.

FIG. 5 shows yet another embodiment 200 wherein a fan rotor 202 and afirst stage compressor 204 rotate at a common speed. The gear reduction206 (which may be structured as the geared architecture 45 disclosedabove) is intermediate the compressor rotor 204 and a shaft 208 which isdriven by a low pressure turbine section.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldunderstand that certain modifications could come within the scope ofthis disclosure. For these reasons, the following claim should bestudied to determine the true scope and content of this disclosure.

What is claimed is:
 1. A gas turbine engine comprising: an engine centerline longitudinal axis; a fan section including a fan with a plurality of fan blades wherein the fan section drives air along a bypass flow path in a bypass duct and the fan rotates about the engine centerline longitudinal axis; a low corrected fan tip speed less than 1400 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °R)/(518.7 °R)]^(0.5), where T represents the ambient temperature in degrees Rankine; a fan pressure ratio less than 1.38 at cruise conditions of 0.8 Mach and about 35,000 feet, wherein the fan pressure ratio is measured across a fan blade alone; a speed reduction device comprising an epicyclic gear system having five intermediate gears and a gear ratio of at least 2.6 and less than or equal to 4.1; a plurality of bearing systems; a low pressure turbine in communication with a first shaft; a high pressure turbine in communication a second shaft; and a mid-turbine frame arranged between the high pressure turbine and the low pressure turbine, wherein the mid-turbine frame supports at least one bearing system and one or more airfoils that extend in a flow path; and wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device; wherein the low pressure turbine includes at least three stages and no more than four stages; wherein the high pressure turbine includes two stages; and wherein the low pressure turbine includes at least one rotor constrained by a first stress level, at least one of the plurality of fan blades of the fan constrained by a second stress level and having a fan tip speed boundary condition, and the gear ratio is configured such that in operation the fan blade does not exceed the fan tip speed boundary condition or the second stress level, and the low pressure turbine rotor does not exceed the first stress level.
 2. The gas turbine engine of claim 1, wherein the low pressure turbine includes four stages.
 3. The gas turbine engine of claim 2, wherein the gear system is a star gear system with a ring gear, and a sun gear, wherein the gear ratio is determined by measuring a diameter of the ring gear and dividing that diameter by the diameter of the sun gear.
 4. The gas turbine engine of claim 2, a bypass ratio greater than 11.0 and less than 22.0 at cruise conditions of 0.8 Mach and about 35,000 feet.
 5. The gas turbine engine of claim 2, a bypass ratio greater than 13 and less than 20 at cruise conditions of 0.8 Mach and about 35,000 feet.
 6. A gas turbine engine comprising: an engine centerline longitudinal axis; a fan section including a fan with a plurality of fan blades wherein the fan section drives air along a bypass flow path in a bypass duct and the fan rotates about the engine centerline longitudinal axis; a low corrected fan tip speed less than 1400 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °R)/(518.7 °R)]^(0.5), where T represents the ambient temperature in degrees Rankine; a fan pressure ratio less than 1.38 at cruise conditions of 0.8 Mach and about 35,000 feet, wherein the fan pressure ratio is measured across a fan blade alone; a speed reduction device comprising an epicyclic gear system having five intermediate gears; a plurality of bearing systems; a low pressure turbine in communication with a first shaft; a high pressure turbine in communication with a second shaft; and a mid-turbine frame arranged between the high pressure turbine and the low pressure turbine, wherein the mid-turbine frame supports at least one bearing system and one or more airfoils that extend in a flow path; and wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device; wherein the high pressure turbine includes two stages; and wherein the low pressure turbine includes at least one rotor constrained by a first stress level, at least one of the plurality of fan blades of the fan constrained by a second stress level and having a fan tip speed boundary condition, and the gear system is configured such that in operation the fan blade does not exceed the fan tip speed boundary condition or the second stress level, and the low pressure turbine rotor does not exceed the first stress level.
 7. The gas turbine engine of claim 6, wherein the low pressure turbine includes five stages.
 8. The gas turbine engine of claim 6, further comprising a low pressure compressor including three stages, wherein the low pressure turbine drives the low pressure compressor.
 9. The gas turbine engine of claim 6, wherein the low pressure turbine includes at least three stages and no more than four stages.
 10. The gas turbine engine of claim 6, wherein the speed reduction device includes a gear ratio less than or equal to 4.1.
 11. The gas turbine engine of claim 10, wherein the gear ratio is greater than or equal to 2.6.
 12. The gas turbine engine of claim 9, wherein the low pressure turbine includes four stages.
 13. The gas turbine engine of claim 12, a bypass ratio greater than 11.0 and less than 22.0.
 14. The gas turbine engine of claim 13, wherein the bypass ratio is at cruise conditions of 0.8 Mach and about 35,000 feet.
 15. The gas turbine engine of claim 12, further comprising a low pressure compressor including three stages, wherein the low pressure turbine drives the low pressure compressor.
 16. The gas turbine engine of claim 15, further comprising a high pressure compressor including eight stages, wherein the high pressure turbine drives the high pressure compressor.
 17. The gas turbine engine of claim 15, wherein the gear system is a star gear system with a ring gear, a sun gear, and a star gear ratio, and the star gear ratio is determined by measuring a diameter of the ring gear and dividing that diameter by the diameter of the sun gear.
 18. The gas turbine engine of claim 17, wherein the gear system includes a carrier supporting the intermediate gears and fixed from rotation relative to a static structure of the gas turbine engine.
 19. The gas turbine engine of claim 6, a bypass ratio greater than 13 and less than 20 at cruise conditions of 0.8 Mach and about 35,000 feet.
 20. A gas turbine engine comprising: an engine centerline longitudinal axis; a fan section including a fan with a plurality of fan blades wherein the fan section drives air along a bypass flow path in a bypass duct and the fan rotates about the engine centerline longitudinal axis; a low corrected fan tip speed less than 1400 ft/sec, and the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °R)/(518.7 °R)]^(0.5), where T represents the ambient temperature in degrees Rankine; a bypass ratio of greater than 11.0 and less than 22 at cruise conditions of 0.8 Mach and about 35,000 feet; a fan pressure ratio less than 1.38 and greater than 1.25 at cruise conditions of 0.8 Mach and about 35,000 feet, wherein the fan pressure ratio is measured across a fan blade alone; a speed reduction device comprising an epicyclic gear system having five intermediate gears; a plurality of bearing systems; a low pressure turbine in communication with a first shaft and includes a pressure ratio greater than about 5:1, the low pressure turbine includes an inlet having an inlet pressure, and an outlet having an outlet pressure, and the pressure ratio of the low pressure turbine is a ratio of the inlet pressure to the outlet pressure; a high pressure turbine in communication with a second shaft; and a mid-turbine frame arranged between the high pressure turbine and the low pressure turbine, wherein the mid-turbine frame supports at least one bearing system and one or more airfoils that extend in a flow path; and wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device; wherein the low pressure turbine includes at least three stages and no more than four stages; and wherein the low pressure turbine includes at least one rotor constrained by a first stress level, at least one of the plurality of fan blades of the fan constrained by a second stress level and having a fan tip speed boundary condition, and the gear system is configured such that in operation the fan blade does not exceed the fan tip speed boundary condition or the second stress level, and the low pressure turbine rotor does not exceed the first stress level.
 21. The gas turbine engine of claim 20, wherein the high pressure turbine includes two stages.
 22. The gas turbine engine of claim 21, wherein the low pressure turbine includes three stages.
 23. The gas turbine engine of claim 21, wherein the gear system has a gear ratio of less than or equal to 4.1.
 24. The gas turbine engine of claim 23, wherein the gear reduction ratio is at least 2.6.
 25. The gas turbine engine of claim 24, wherein the gear system is a star gear system with a ring gear, and a sun gear, wherein the gear ratio is determined by measuring a diameter of the ring gear and dividing that diameter by the diameter of the sun gear.
 26. The gas turbine engine of claim 24, wherein the low pressure turbine includes four stages.
 27. A gas turbine engine comprising: an engine centerline longitudinal axis; a fan section including a fan with a plurality of fan blades wherein the fan section drives air along a bypass flow path in a bypass duct and the fan rotates about the engine centerline longitudinal axis; a low corrected fan tip speed less than 1400 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °R)/(518.7 °R)]^(0.5), where T represents the ambient temperature in degrees Rankine; a fan pressure ratio less than 1.38 and greater than 1.25 at cruise conditions of 0.8 Mach and about 35,000 feet, wherein the fan pressure ratio is measured across a fan blade alone; a speed reduction device comprising an epicyclic gear system having five intermediate gears and a gear ratio of at least 2.6 and less than or equal to 4.1; a plurality of bearing systems; a low pressure turbine in communication with a first shaft; a high pressure turbine in communication with a second shaft; a mid-turbine frame arranged between the high pressure turbine and the low pressure turbine, wherein the mid-turbine frame supports at least one bearing system and one or more airfoils that extend in a flow path; and wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device; wherein the low pressure turbine includes four stages; and wherein the low pressure turbine includes at least one rotor constrained by a first stress level, at least one of the plurality of fan blades of the fan constrained by a second stress level and having a fan tip speed boundary condition, and the gear ratio is configured such that in operation the fan blade does not exceed the fan tip speed boundary condition or the second stress level, and the low pressure turbine rotor does not exceed the first stress level.
 28. The gas turbine engine of claim 27, wherein the high pressure turbine includes two stages.
 29. The gas turbine engine of claim 27, a bypass ratio greater than 11 and less than 22 at cruise conditions of 0.8 Mach and about 35,000 feet.
 30. The gas turbine engine of claim 29, wherein the low pressure turbine has a pressure ratio greater than 5:1, the low pressure turbine includes an inlet having an inlet pressure, and an outlet having an outlet pressure, and the pressure ratio of the low pressure turbine is a ratio of the inlet pressure to the outlet pressure. 